Composite structures provide several advantages over metallic structures. For example, composite structures can be configured to provide high specific stiffness and high specific strength relative to metallic structures. Furthermore, composite structures can be tailored to provide a relatively high degree of strength and stiffness along a primary load path. The ability to tailor the strength and stiffness of composites may result in lightweight structures. In addition, composite materials may have improved fatigue resistance relative to metallic materials and are more resistant to corrosion.
Composite structures may be formed as a stack of relatively thin layers or plies that are laminated together. Each ply in the composite laminate may include fibers that serve as the primary load-carrying constituent. The composite material may be formed as unidirectional tape wherein the fibers in each ply are oriented parallel to one another and are held in position by a matrix constituent such as an epoxy resin. The matrix constituent may also redistribute loads between adjacent fibers.
The composite structure may be configured such that the fibers in one ply are oriented in the same direction of the composite structure or in a different direction than the fibers in adjacent plies. The relative orientations of the plies may be selected to provide the desired strength and stiffness characteristics of the composite structure. Each ply in the composite laminate may be formed of the same material system. However, composite structures may also be formed as hybrid structures containing plies formed of different materials to achieve a desired design objective. For example, a hybrid composite structure may primarily include plies formed as unidirectional carbon fiber tape for load-carrying purposes. The hybrid composite structure may also include one or more outer plies formed of woven fiberglass cloth to provide improved impact resistance to the composite structure.
Conventional methods of designing a composite structure include constructing a finite element model (FEM) of the structure and subjecting the FEM to loads to determine the stresses and strains in the structure and to perform sizing of the structure to meet strength, stiffness and weight requirements. An FEM is typically comprised of a mesh of multiple finite elements. Each element may represent one or more components or sub-components of the composite structure. For example, a barrel section of an aircraft fuselage may include a composite laminate skin attached to a series of frames and stringers, each of which may also be formed as composite laminates. An FEM of the fuselage may be constructed simulating the skin and stringer geometry. The composite laminate stringer may be formed in a closed hat shape made up of several sub-components such as a cap, a pair of webs, a pair of flanges, an inner wrap laminate, and a plank laminate which may be bonded or co-cured together after which the skin may be bonded to the stringers and other structural components that make up the barrel section.
The process of designing the barrel section of the fuselage may include optimizing several design variables. Such design variables may include the geometry of the components and subcomponents that make up the fuselage. The geometry may include the size (i.e., length, width, height) and the shape of the components and sub-components. For example, the geometry design variables for a stringer may include the width of the cap, the width of the flanges, the height of the webs above the flanges, the angle at which the webs are oriented relative to the flanges, and other geometry design variables.
Additional design variables that may be optimized in the design process include the ply arrangement for the composite laminates of the skin and the stringer to meet strength, stiffness, weight and other requirements. Conventional methods of designing composite laminates include a determination of the stacking sequence of the laminate including a determination of the individual ply thickness, the fiber angle of each ply, and the relative location of the ply in the through-thickness direction. For certain structures, loading conditions may dictate a laminate thickness requiring a relatively large quantity of plies. For example, a wing panel of an aircraft may require up to one hundred or more plies of composite material, each of which requires the determination of the fiber angle and the ply thickness. As may be appreciated, a ply-by-ply determination of such a stacking sequence for relatively thick composite laminates adds a large quantity of design variables to the design process which significantly increase the complexity of the design process.
A further design variable that may be included in designing a composite structure is the material system of the plies that make up the composite laminate. For example, as indicated above, it may be desirable to form the composite structure as a hybrid comprised of two or more material systems. For example, a hybrid composite structure may include plies of unidirectional composite tape for load-carrying purposes and one or more outer plies of composite cloth for impact resistance or other purposes. Unfortunately, conventional design methods are not understood to provide an efficient means for optimizing hybrid composite laminate containing two or more material systems.
Even further, conventional methods of designing composites plates such as Classical Lamination Theory do not account for transverse shear deformation which may be an important consideration in relatively thick composite laminates. Transverse shear deformation may occur when a structure is subjected to certain loads or combinations of loads that result in shear stresses and shear strains contributing to shear deformation in the structure. Under relatively high shear loads, excessive transverse shear deformation may compromise the buckling stability of the composite laminate. Failure to account for transverse shear deformation in relatively thick composite laminates may result in under-conservative designs.
As can be seen, there exists a need in the art for a system and method for optimizing a composite structure that can accommodate a large quantity of plies in a computationally efficient manner and which may account for transverse shear deformation. Furthermore, there exists a need in the art for a system and method for optimizing hybrid composite laminates having composite plies formed of two or more material systems.